Internal-combustion geared turbine



March 9, 1954 c. F. RocHl-:vlLLE 2,671,315

INTERNAL-COMBUSTION GEARED TURBINE Filed Nov. l2, 1948 3 Sheets-Sheet lIIIIII IN VEN TOR C/)a e: F.l FPocbevi/k ATrR/IE Ys ,E willlllllmm,

March 9, 1954 Q F; ROCHEWLLE 2,671,315

INTERNAL-COMBUSTION GERED TURBINE Filed Nov. l2, 1948 5 Sheets-Sheet 3IN VEN TOR. Char/es F.' Fac/7e vil/e Patented Mar. 9, 1954 UNITED STATESFATENT OFFlCE INTERNAL-COMBUSTION GEARED TURBINE Application Novemberl2, 1948, Serial No. 59,496

7 Claims.

This invention is directed to internal combustion turbine engines foraircraft, and particularly refers to an improved arrangement in which aplurality of fuel burners and combustion cham bers are disposed in arotor with their longitudinal axes substantially in the plane ofrotation at right angles to the axis of the turbine, and are arcuatelypositioned in helically overlapping relation to discharge combustionproducts tangentially to indu-ce rotation by reaction, the rotor beingpositively geared to counter-rotating varied turbine wheels whichreceive the combustion products to be rotated thereby, both rotatingelements being adapted to drive a propeller shaft. Desirably, but notnecessarily, clutch provisions and free-wheeling means for driving aircompressing means may be included.

It is an object of this invention to provide a construction of thenature set forth which will permit high relative speed of the combustiongas driven counter-rotating elements, with a substantially lower speedof the propeller drive shaft.

Another object is to provide an improved arrangement of the burner andcombustion chambers of a turbine of this type, to give a relatively longpath for combustion to take place, in a structure of relatively shortaxial length.

Another object is to provide an improved arrangement of stationaryfuel-air nozzles around the periphery of the turbine, said nozzlesfeeding continuously into a heat exchange passage and thence into aperforated burner tube to initiate combustion, which is subsequentlysubstantially completed in a combustion chamber.

Another object is to provide an improved ar# rangement of gearing andlubricating means for a turbine unit of this type.

Another object is to provide an improved control means for the aircompressing means of a turbine of this type, that can be selectivelyconnected to furnish air for combustion at air speeds below that whichram-jet action occurs, and may be disconnected from the power unit atspeeds where air compression is not required to `consume the fuel at thedesired rate.

These and other objects and advantages of the invention will 'be furtherapparent from the following description and the attached drawings, whichform a part of this specification, and illustrate a preferred embodimentas applied to an aircraft engine.

(Cl. Gli-39.35)

In the drawings, Figure 1 is a longitudinal part sectional view of theupper half of the forward or compressor section of the engine of thisexample, and illustrates a preferred arrangement of starter, clutch andfree-wheeling means. Figure 2 is a rearward continuation view of thecentral part of the arrangement of Figure l, also in longitudinal andsubstantially complete cross section, illustrating the gearing,combustion chambers and turbine elements. Figure 3 is fa diagrammaticand somewhat simplified trans-- verse sectional view taken on lineIII-II of Figure 2, looking forward and taken substantially in the planeof the longitudinal axis 'of one of the two combustion chambers, showingthe heli'- cally overlapping arrangement of the said ychl'ain'- bers.

Referring to Figure l, reference numeral i0 designates -a generallycylindrical metal outer casing for the engine unit, which 'may be fairedinto the wing structure ii of an airplane. An inner metal housing l2 is"suitably spaced and supported from the outer casing' and forms anannular passage I3, open at its forward end 'to admit cooling air forthe housing and some engine parts to be described below. Extendinginwardly from inner housing l2 are a plurality of struts l5 to supportbearing I6 for the fori ward end of propeller shaft Il, the latterdesira-v 'bly being hollow as at I8 to admit cooling air. A conventionalpropeller spinner or hub 'I9 supporting propeller blades 20 is secured'to the outer end of shaft Il.

A generally conical air guide or ram-jet cone 2i extends from theforward end of bearing iB to a set of stationary guide vanes 22 at theinlet of a multi-stage air `compressor formed by a plurality of rows ofinwardly directed stationary vanes 23 and outwardly directed rotatingvanes 2li, the latter supported on compressor rotor 25 turning onbearings 26 and adapted to be selectively driven from shaft ll throughclutch 28 and free-wheeling means 29 which may be a conventionalarrangement of spring-controlled balls on inclined wedges, so arrangedthat when more power is required than for idling or gliding, theincreased speed of rotation of Vthe 'gas actuated power elements to bedescribed below will engage the free-wheeling means and 4drive thecompres'i scr rotor to give the desired air now to the com-I bustionchamber. Upon reaching air speeds in the neighborhood of 400 M. P. H.,the ram-jet action of the air intake between i2 and 2| will be adequatefor air supply, whereupon clutch 28 may be disengaged, leaving the aircompressor rotor to idle or windmill. In this example a starting motor29| is provided to drive the propeller shaft I1 through the medium ofgears 30 and 3I. 'I'he clutch control means generally designated 32 maybe of the conventional hydraulic type.

Referring to the main section of the turbine engine, and particularly toFigure 2, there is illustrated the after end of compressor rotor 25,supported at this point on shaft I1 by bearing 26. A stationary spider33 is secured within housing I2 and supports a plurality ofcircumferentially spaced air-fuel mixing nozzles 34, supplied withliquid fuel such as gasoline, kerosene or fuel oil through a conduit orpassage 35, which may surround housing I2, and be connected to anydesired type of pressure feed system, not shown. The forward face ofspider 33 and the after face of compressor rotor are desirably providedwith sealing means, such as the labyrinth of interengaging grooves andlandsI 3B, to prevent undue leakage of fluids at this point.

Stationary spider 33 also serves to support bearings 31 for a pluralityof circumferentially spaced hollow shafts 38, each shaft supporting atits forward end a first planet geai` 33, and supporting at its after enda second planet gear 4B. At the rear end of propeller shaft i1 is keyeda sun gear 4I, adapted to engage and be driven by the forward planetgear 39. On the rearinost portion of shaft I1 is supported a bearing t2for the forward extension of a second sun gear 43, the latter also beingkeyed to the forward end of a hollow turbine shaft 44, which extendsrearwardly to a flange 45, onto the rim of which turbine wheels 46 and41 are secured by a plurality of In this manner, there is provided aturbine shaft 44, supported at its forward end by bearing 42 which ismounted on the rearmost end of propeller shaft I1, through the axiallyelongated annular sun gear 43, and supported at its after end by bearing50, through the medium of the turbine wheels 46 and 41 and flangedmembers 45 and 49. A supplemental bearing for the after end of propellershaft I1 may be installed at 54, this bearing being supported by a platesecured to the forward face of the stationary spider 33 by through bolts56, and also acting as a support for the forward ends of shafts 38through the medium of bearings 51.

A plurality of bearings 58 on turbine shaft 44 serves to support agenerally cylindrical hollow combustion chamber rotor 59, which ispositioned between the after face of stationary spider 33 and theforward face of turbine wheel 46. An annular flange 60 projectsforwardly from the front of rotor 59, and is provided with internal gearteeth to engage the after planet gears 49. This arrangement, whereby sungear 43 on turbine shaft 44 engages planet gears 49 at their inner pitchcircle, and the teeth of annular flange gear 60 engages the same gears40 at their outer pitch circle, insures that the turbine shaft 44 andthe turbine wheels 46 and 41 carried thereby, will rotate in theopposite direction from the combustion chamber rotor 59, for reasonswhich will b-e further apparent below. Desirably, the forward face ofcombustion chamber rotor 59 is sealed from stationary spider 33 by meanssuch as the grooved labyrinth 6I, and the after face of rotor 59 issealed from the turbine wheel 4S by means such as the interengaginggrooves and lands 32.

'Ihe forward face of rotor 59 is provided adjacent its periphery with anannular opening 63, extending substantially throughout itscircumference, whereby the fuel-air mixture from the nozzles 34 mayenter the interior passages G4 and 65 of the rotor. These passages areformed by symmetrical annular metal housings 83 supported within therotor by ported rings 51, the housings being generally rectangular incrosssection, with an enlarged entrance scoop (Figure 3) and, inaddition, similarly ported as at 63 (Figure 2) to permit the fuel-aircombustible mixture, after it is heated by the means to be describedbelow, to pass into the spaces-59 within the housings and thence intothe circumferentially arranged burner tubes 10 (Figure 3) which lie inhelically overlapping relation within the housings 65 and rotor 59. Inthis example, only two of such housings and tubes are illustrated. andthe transverse section of rotor 59 in Figure 3, which illustratesdiagrammatically how they are disposed, has been somewhat simplified tofacilitate this description. Obviously, if additional burners aredesired, they would be symmetrically arranged in similar helicallyoverlapping relation Within the rotor. Desirably, but not necessarily,the inlets and outlets of rotor 59 are at the same distance from theaxis of its rotation, thus reducing the diameter of the unit, andfacilitating its construction and maintenance.

Referring to Figure 3, it will be noted that the fuel-air mixture fromspace 69 enters the burner tubes '.'0 through the ports 1I, to beignited by means such as the spark plug 12, the latter supplied by highvoltage electric current by means which will be described below. Afterthe combustion is initiated in burner tube 19, the heated gases passinto the combustion chamber or space designated 13, where expansiontakes place to give a high velocity to the gases, which are thendirected out of the discharge end of the combustion space by vanes 14 toimpinge against the blades 15 of the rst wheel 46 or stage of theturbine element of the unit (Figure 2). After leaving that stage, thegases are straightened by stationary vanes 16, the latter supported bythe after extension 11 of housing I2, and are directed into the vanes 18of the second wheel 41 or stage of the turbine element of the engineunit. In this example, where only two such stages are illustrated, theexhaust gas, from which the greater part of the energy has beenabstracted by the turbine vanes, emerges into the exhaust passage 19formed between the housing extension 's1 and exhaust cone 52, to bereleased to the atmosphere, where it may impart added energy to theairplane by its reaction as a jet.

The helically overlapping arrangement of the housings 66, burner tubes10 and combustion chambers 13, with the surrounding passages 64 and 65,insure a maximum heat transfer' from these highly heated surfaces to theincoming fuelair mixture, thus servingto cool the parts below thosetemperatures at which damage would occur and also adding heat energy tothe combustible mixture and increasing the ,overall efdciency of theturbine engine. From Figure 2 it is. apparent that the last-namedmixture divides, as it enters the rotor, part of it going into passagesdfi and the remainder of it into passages 65 to. circulate within therotor as indicated by the arrows and nally to merge again as it passesthrough ports 68 into spaces 69 inside neusingst. From those spaces thenow-heated fuel-air mixture enters theY burner tubes 'i8 of the nextcircumferentially adjacent housings 66, as best shown schematically inFigure 3. The materials of these partsl are desirably of heat-resistantalloy, such as an alloy of chromium, iron and molybdenum.

After the turbine engine is once in operation, it will not ordinarily benecessary to continue the supply of high tension electric current tospark plugs l2, but foi` starting, and for reignition the current istransmitted from any suitable source through a brush 8d (Figure 2) to aninsulated annular metal collector ring 8l on the forward face ofcombustion chamber rotor 59. An insulated lead 82 extends from ring 8ito the terminal of each spark plug l2.

As stated above, the propeller shaft El is desirably hollow throughoutits length. In this example a sleeve 84 is mounted within the turbineshaft flri, and extends within but is spaced from the bore IB ofpropeller shaft il. rlhis not only serves to conduct cooling air intothe turbine shaft, to cool its bearings 553 and 58, but also suppliescooling air to the turbine wheels 46 and 4l, which may be ported as at85 and 35, respectively, to admit air to the hollow vanes or blades 'i5and i8, from which it escapes into exhaust passage i9 through notches 8land 88 in the trailing edges of the blades. Sleeve 315 also serves toconduct lubricant from the gear and bearing space within the stationaryspider 33 through the helical passage formed by blades 89 to bearings58.

In those applications of this invention which are not carried out in astructure, such as an airplane, which moves relatively to the air usedfor combustion, the clutch means 23 would ordinarily not be required asthe compressor rotor 25 would operate at all times that the power shaftIl is rotated. Similarly, circumstances might also be encounteredwherein the freewheeling means 29 could be omitted, as will be obviousto those skilled in the art.

In conclusion, it will be appreciated from the foregoing descriptionthat this invention comprehends broadly the combination, in an internalcombustion turbine engine, of burner tubes and combustion chambersoperating by the reaction of gases and discharging into vaned turbineelements, said chambers and elements being positively and compactlygeared to rotate in opposite directions and to jointly drive a singleshaft at a lower speed than that at which either rotates to deliveruseful power, for example, to an airplane propeller. It includes alsothe circumferential helically overlapping arrangement for the housings,burner tubes and combustion chambers with preheating and heat transferpassages for the fuel-air mixture. Although a single embodiment isdescribed and illustrated as applied to an aircraft engine, it isobvious that numerous modifications and changes could be made in itsconstruction and application to other uses without departing from theessential features of the invention, and all such as fall within thescope of the appended rotating u? claims are understood: to be embracedthereby.

I" claim:

1. In combination in an internal combustion turbine engine, a rotorSupporting a plurality of combustion chambers, means for supplying acombustible mixture to said chambers, at least one turbine wheelforreceiving products of combustion from al1- oi'saidv chambers, a powershaft coaxial; with said rotor, and direction reversing gear mea-ns'separately connecting said shaft to said rotor and said turbine wheel torotate' them in Vopposite directions to deliver power to said 2. Acombination according to claim -1,= in which said combustion chambersare arcuately disposed in helically overlapping relation around saidrotor.

3. A combination according to claim 1, with the addition of a rotatableair compressor means surrounding said shaft, and means for selectivelyconnecting said compressor means to said shaft.

4. In combination in an internal combustion turbine engine, a pluralityof circumferentially arranged fuel-air nozzles, axial flow means forIchambers in said rotor, a vaned turbine wheel for continuouslyreceiving heated gases from said combustion chambers, a power shaft, andgear means connecting said rotor and said turbine wheel to rotate inopposite directions, and to rotate said power shaft.

5. A combination according to claim 4, with the addition of a, housingin said rotor for said tubes and chambers forming passages to conductsaid combustible mixture in heat exchange relation thereto prior to itsintroduction into said burner tubes.

6. In combination in an internal combustion turbine engine having anannular air passage and a plurality of fuel-air mixing nozzles at theend of said passage, a rotor adjacent the end of said passage having anannular inlet for combustible mixtures in one end face, a plurality ofheat exchange housings in said rotor in helically overlapping relation,a perforated burner tube communicating with and contained in each ofsaid housings, ignition means in each of said tubes, a combustionpassage leading from each of said tubes, means for passing saidcombustible mixture from said annular inlet first through said housings,around said combustion passages to be heated thereby, and then into saidburner tubes to be ignited therein, and a combustion gas outlet for eachsaid passage in the opposite end face of said rotor.

'7. A combustible mixture igniting and consuming means for an internalcombustion engine having a turbine element to receive the products ofcombustion from said means, comprising a rotor having opposed axiallyspaced faces,

an annular passage formed in one of said faces, a plurality ofcircularly disposed helically overlapping housings in said rotor forcontinuously receiving a combustible mixture from said passage, aperforated burner tube in each of said housings, an enclosed combustionchamber for each of said burner tubes, an outlet for each of -saidcombustion chambers, each of said housings adapted continuously toconduct a portion of said combustible mixture in heat exchange relationwith its associated combustion chamber and Ydirect said portion of saidcombustible mixture to a succeeding housing and burner tube to beignited therein.

CHARLES F. ROCHEVILLE.

References Cited in the le of this patent UNITED STATES PATENTS NumberName Date Clark Mar. 6, 1906 Kothe Oct. 6, 1914 Holtz July 13, 1915 LakeSept. 17, 1918 Tyler Jan. 14, 1919 Number 10 Number

